Rocket motor and method

ABSTRACT

7. The method of operating a solid-fueled rocket motor to vary the total impulse achieved, said motor comprising, in combination, a substantially cylindrical outside wall which forms a chamber, said wall being tapered at the rearward end of said chamber to form a converging-diverging nozzle, said wall also closing the forward end of said chamber except for an opening to receive a tube whereby fluid can be introduced into said chamber; a solid rocket propellant grain rigidly disposed in said chamber, said grain comprising a plurality of concentric layers of propellant, the propellant layer nearest the center of the grain having disposed therein a perforation passing from end to end therethrough along the center axis of said grain, said layers consisting essentially of a cured intimate mixture of polymeric binder, fuel particles, and an oxidizer therefor, said layers being completely separated from each other by a liner, said liner consisting essentially of a cured polymeric material incapable of supporting its own combustion, said method comprising: A. IGNITING AND BURNING THE INNERMOST LAYER OF PROPELLANT, THEREBY EXPOSING THE ADJACENT LINER; B. INTRODUCING INTO THE FORWARD END OF SAID CHAMBER THROUGH SAID OPENING AN OXIDIZER FLUID WHICH IS HYPERGOLIC WITH SAID LINER WHEREBY THE LINER IGNITES, BURNS, AND IN TURN IGNITES THE ADJACENT LAYER OF PROPELLANT; AND C. REPEATING STEPS (A) AND (B) UNTIL AS MANY PROPELLANT LAYERS AS DESIRED HAVE BEEN CONSUMED.

United States Patent Fink et a1.

[ 51 Ju1y18,1972

[ ROCKET MOTOR AND NETHOD [72] Inventors: Robert H. Fink; Eugene ,1.Palm, both of Huntsville, Ala.

[73] Assignee: The United States of America as represented by theSecretary of the Army [22] Filed: March 11, 1964 [21] Appl No.: 351,256

[52] US. Cl ..60/220, 60/207, 60/250, 60/219, 60/253, 102/103, 149/2,149/19 [51] Int. Cl. ....C06d 5/10, F02k 9/06 [58] Field of Search149/2, 17-19; 102/98, 103; 60/35.4, 207, 219, 220, 250, 251, 253

3,032,970 1/1957 Fox.. ..60/253 Primary Examiner-Benjamin R. PadgettAttorney-Harry M. Saragovitz, Edward J. Kelly and Herbert Berl EXEMPLARYCLAIM 7. The method of operating a solid-fueled rocket motor to vary thetotal impulse achieved, said motor comprising, in combination, asubstantially cylindrical outside wall which forms a chamber, said wallbeing tapered at the rearward end of said chamber to form aconverging-diverging nozzle, said wall also closing the forward end ofsaid chamber except for an opening to receive a tube whereby fluid canbe introduced into said chamber; a solid rocket propellant grain rigidlydisposed in said chamber, said grain comprising a plurality ofconcentric layers of propellant, the propellant layer nearest the centerof the grain having disposed therein a perforation passing from end toend therethrough along the center axis of said grain, said layersconsisting essentially of a cured intimate mixture of polymeric binder,fuel particles, and an oxidizer therefor, said layers being completelyseparated from each other by a liner, said liner consisting essentiallyof a cured polymeric material incapable of supporting its owncombustion, said method comprising:

a. igniting and burning the innermost layer of propellant,

thereby exposing the adjacent liner;

b. introducing into the forward end of said chamber through said openingan oxidizer fluid which is hypergolic with said liner whereby the linerignites, burns, and in turn ignites the adjacent layer of propellant;and

c. repeating steps (a) and (b) until as many propellant layers asdesired have been consumed.

10 Claims, 3 Drawing Figures PATENTED JUU 8 m2 ROCKET MOTOR AND METHODThe invention described herein may be manufactured and used by or forthe Government for governmental purposes without the payment of anyroyalty thereon.

This invention relates to solid-fueled rocket motors. Particularly theinvention concerns a novel solid rocket propellant grain structure, amethod for using these grains in the operation of a solid-fueled rocketengine in such a manner as to regulate the overall impulse of theengine, and to solid rocket motor themselves.

Solid-fueled rocket motors possess certain characteristics which renderthem preferable to liquid-fueled engines, especially for militaryapplication. Thus, solid-fueled rocket engines do not requirecomplicated systems of valves, pumps, and tubing to move the propellantconstituents from their storage tanks to the combustion chamber.Moreover, there is no requirement for the storage, transportation, andhandling of great quantities of highly corrosive, toxic, and easilyvaporized materials as with many liquid-fueled engines.

However, liquid-fueled engines have one very important and distinctadvantage over the solid-fueled engines. In a solid propellant grain,the fuel particles and oxidizer are bound together as a single unit.Once the grain has been ignited, combustion continues until the entiregrain is consumed. This makes it difiicult to vary the total impulse ofthe rocket motor. Such devices as head-in reversal ports have not provencompletely satisfactory in alleviating this problem. On the other hand,the flow of liquid propellant to the combustion chamber can be readilycontrolled by valves, variable pump speeds, and the like so that theoverall total impulse of the engine can be varied. The flow ofpropellant can even be completely stopped and then restarted if desired.This characteristic of liquid-fueled engines permits propelling missilesvarying distances. For example, a liquid-fueled artillery rocketdesigned to deliver a warhead up to 250 miles can also deliver thewarhead to any reasonable point less than 250 miles simply bycontrolling the flow rate and total flow of liquid propellant to thecombustion chamber. A solid-fueled engine designed to deliver a warheada specific distance in free flight is limited in its ability to alterthis range capability by means other than changing the angle ofelevation of the rocket at launch time, which is far from being acompletely satisfactory solution to the problem.

It is readily apparent that it would be very desirable to have asolid-fueled rocket motor that would permit varying to total impulse,thus combining the advantages of both the solidfueled and liquid-fueledengines.

According to the present invention, the solid rocket propellant grainsconsist of a plurality of individual layers of propellant. The grainsare cast in concentric layers with each layer of propellant beingcompletely separated by a liner made of material which does not have any(or at least insuficient) oxidizer incorporated therein to sustain itsown combustion. When the innermost layer of propellant is ignited itburns through completely. However, the liner material separating thepropellant layers does not burn as there is no available oxidizer tosupport its combustion. Therefore, the remaining layers of propellant donot ignite and the rocket engine ceases to function. If, however, theintended range of the rocket requires additional impulse, an oxidizerfluid which is hypergolic with the liner is injected into the chamberand thereby ignites the liner which, after burning through, ignites theadjacent layer of propellant. This process is repeated until asufficient number of propellant layers has been burned to provide thenecessary impulse to propel the payload of the rocket the desireddistance. The instant invention makes it possible to design solid-fueledrocket engines which can be used to propel a payload varying distancesby regulating the amount of the solid propellant grain consumed. Bycombining the regulation of the amount of fuel burned with a variationin launch angle, the instant invention permits the use of only one typeof solidfueled rocket motor to provide a wide range of performancecapabilities with regard to effective range and trajectory. Such anengine combines the simplicity and reliability of the solid fueledengines with the variable range and trajectory capabilities of theliquid-fueled engines. Moreover, an engine of the type contemplated bythe invention offers tremendous economic advantages since the many typesof solid-fueled engines now required to give complete range capabilitiesover a wide latitude of operating distances could be replaced with afewer number of engines of a type offering a wider distribution of rangeand trajectory capabilities. Furthermore, many liquid-fueled engineswould no longer be necessary and this would eliminate the costlyconstruction of engines of this type as well as the time-consuming andexpensive handling and storage problems associated with liquid-fueledengines, especially those which utilize highly corrosive and/or toxicpropellants.

In accordance with the foregoing, it is an object of the presentinvention to provide a solid rocket propellant grain of uniquestructure.

Another object of the invention is to provide a solid rocket propellantgrain which, when burned in a rocket engine, can deliver predeterminedlevels of impulse.

A still further object of the instant invention is to provide a methodfor operating a solid-fueled rocket engine whereby total impulses ofvarying predetermined magnitude can be accomplished.

An additional object of the invention is to provide a solidfueled rocketmotor capable of producing a varying total impulse.

The manner in which these and other objects can be accomplished willbecome apparent from the following detailed description wherein:

FIG. 1 is a view partially in section of one embodiment of a solidrocket motor and a propellant grain contemplated by the invention;

FIG. 2 is a sectional view of the propellant grain and the rocket motortaken along line 2-2 of FIG. 1; and

FIG. 3 is a sectional view of another embodiment of a propellant graincontemplated by the present invention.

It will be apparent that the present invention is independent of theparticular propellant composition utilized in fabricating the grains.The propellant composition will normally be a cured intimate mixture ofa polymeric binder, fuel particles, and an oxidizer. As polymericbinders the polyurethanes, polydienes, polysulfides, polylactams, andthe like are completely satisfactory. Moreover, double-base propellantcompositions are completely satisfactory. If desired, the propellantgrain can contain no fuel components other than the binder. However,high energy fuels can, and normally will, be incorporated in thecomposition. Illustrative of these highenergy fuels are powdered metaland metal hydrides as exemplified by aluminum, aluminum hydride,magnesium, lithium, lithium hydride, boron, and boron hydride and alloysof these metals. As an oxidizer, inorganic nitrates, chlorates,perchlorates, and the like such as ammonium perchlorates, ammoniumnitrate, alkali and alkaline earth metal chlorates, perchlorates, andnitrates (e.g. sodium nitrate, potassium nitrate, potassium perchlorate)will normally be employed. Typical illustrative propellant formulationsand the methods for preparing them are given in the following U.S. Pat.Nos. 2,962,368; 2,982,638; 2,995,430; 2,997,375; 2,997,376; 3,003,861;3,036,939.

The liner which separates the layers of propellant composition can beany material which will not support its own combustion or readilydisintegrate as a result of the heat and erosion caused by the burningof adjacent propellant layers. In addition to these characteristics, theliner material must be hypergolic with some fluid oxidizer such asfluorine, chlorine trifluoride, perhalogenyl fluoride, perhalogenylfluoridehalogen fluoride mixtures, perhalogenyl fluoride-fluorinemixtures, perhalogenyl fluoride-chloryl fluoride mixtures, chlorinepentafluoride, red fuming nitric acid, and the like. This lastrequirement is not particularly limiting since almost any organicmaterial is hypergolic with these oxidizers, especially red fumingnitric acid, chlorine trifluoride, and mixtures of perchloryl fluorideand at least one of the group chlorine trifluoride, fluorine, or chlorylfluoride. For this reason the liner materials will generally be a curedorganic polymer.

The type of organic polymer is not critical and any of the polymericmaterials used as binders and/or fuels in solid propellant compositionscan be used as the liner material to separate the propellant layers. Thepolymeric hydrocarbons and their halogenated analogs are particularlywell suited as liners. In this group would be the l,4-addition polymersand copolymers of the conjugated dienes (polydienes) of up to eightcarbon atoms and their halo substituted analogs as exemplified bypolybutadiene, polychloroprene, polyisoprene, and copolymers of theseconjugated dienes; polyethylene; polybutylene, polystyrene, andpolypropylene. The polysulfides such as those sold by Thiokol ChemicalCorporation as LP- 2, LP- 3, and LP- 8 are useful as liners. Thepolyepoxides, polyurethanes, polylactams constitute another class ofpolymers well suited for use as liners.

From the foregoing, it is apparent that the function of the liner is toseparate the propellant layers from each other so that combustion of onelayer does not automatically lead to the combustion of the adjacentlayer. Therefore, the liner should completely cover the common surfacesof adjacent propellant layers. By common surface is meant those surfaceswhich would be in contact if no liner was disposed between the layers.

The propellant grains of the present invention are substantiallycylindrical and have an internal perforation running from end to endtherethrough along the center axis. Ordinarily the grain will have fromtwo to about eight concentric layers of propellant extending outwardfrom the center of the grain, each layer of propellant being separatedfrom adjacent layers by a coating of the liner material. Obviously,there is no limit on the number of propellant layers a grain may havealthough, in practical application, situations requiring more than eightlayers would be rare. This is especially true since the thickness of thepropellant layers themselves is also easily controlled, thus permittingadditional control over the total impulse of the engine by varying theamounts of propellant in the different layers of propellant.

Any conventional shape can be used for the internal perforation in thegrain including a tubular passage, a fourpointed star, a five-pointstar, and the like. The particular engine and propellant will determinethe best perforation shape for any given application of the invention.

The thickness of each propellant layer can vary over a wide range, forexample, up to 20 inches for ICBMs. However, layer thicknesses of aboutone inch to about 6 inches are most desirable since they are easier tomanufacture. The liner material separating the propellant layers can beup to about 0.75 inch in thickness, but normally will be about 0.1 toabout 0.25 inch in thickness since the thicker the liner, the moreoxidizer required to burn the liner.

The propellant grains are fabricated according to conventionaltechniques. For example, a polysulfide propellant formulation of thetype shown in U.S. Pat. No. 2,997,376 is prepared and poured into a moldhaving a mandrel positioned in the center thereof. After curing, themandrel is removed. Upon removal from the mold, the outside surface ofthe grain is coated with a polymeric liner such as polybutadiene. Theliner can be applied by dipping, spraying, painting, or otherconventional techniques. After each application the liner is allowed tocure or at least to set. Additional liner material is applied until thedesired thickness is obtained. The coated propellant grain thus formedis centered in a mold or motor casing and another layer of propellantformulation is poured around it between the liner and the inside surfaceof the mold or motor casing. The propellant formulation is thenpermitted to cure. in this manner a perforated grain consisting of twoconcentric layers of propellant separated by a cured polymer liner isproduced. Obviously, the procedure can be repeated to produce a grainconsisting of as many propellant layers as desired. Moreover, it shouldbe pointed out that the procedure can be conducted starting with theoutermost propellant layer. Using this technique, a propellantformulation would be poured in a rocket motor chamber or a mold with alarge mandrel centered therein. After the formulation cured, the mandrelwould be removed and a liner of the desired thickness applied to theinside surface of the grain. When the liner had cured, a smaller mandrelwould be centered in the motor casing or mold and additional propellantformulation poured between the liner and the mandrel. This procedurealso produces a perforated grain consisting of two concentric layers ofpropellant separated by a liner material. Again, the procedure can berepeated until the desired number of separated propellant layers havebeen fabricated.

Another method by which the propellant grains of the instant inventioncan be prepared is that of first producing the liners. This can be doneby molding, extruding, or otherwise shaping thin-walled cylinders fromthe above-mentioned liner materials. These preformed liners are thenplaced in suitable molds or the motor cases themselves. Thereafter thepropellant formulation is poured between the liner and the walls of themold or motor case and allowed to cure. Then a second preformed liner ofa smaller size than the first liner is centered inside the first linerand additional propellant formulation poured into the space between thefirst and second preformed liners. The process can be repeated until asmany propellant layers as desired have been achieved. One advantage inthis particular method of preparing the propellant grains of thisinvention is that as many preformed liners as necessary can be placedinside the motor casing or mold and then all the spaces between thevarious liner can be filled simultaneously with the desired propellantformulation. There is also a drawback to this procedure. The preformedliner may not ha ve sufficient strength to support the weight of theuncured propellant without being distorted. In this case, it willnormally be necessary to provide support for the preformed liner untilthe propellant sets or preferably until it cures. This can be done byinserting a support into the inner space of the liner while thepropellant formulation cures. After curing the support is removed andthe process can be repeated. It is obvious that the propellant grainscan be fabricated with preformed liners either from the center positionof the grain outward or from the outermost propellant layer inwardtowards the center of the grain.

The propellant layers and the liners separating them should adhere toeach other strongly so that they are in rigid relationship to eachother. In this manner, the propellant grain acts as a single unit.Achieving this rigid relationship is no problem since the propellantlayers using polymeric binders and polymeric liner materials normallyadhere to each other with great strength. In any situation where theparticular liner and propellant do not exhibit a mutual adherence, athin coating of an adhesive (an epoxy adhesive, for example) or someother material to which both the propellant layer and the liner willadhere is applied to either the propellant layer surface adjacent to theliner or the surface of the liner (or both) thus forming a rigid bondbetween the propellant layer and liner.

In the operation of a rocket engine using the above described propellantgrains, the innermost layer of propellant is ignited in the conventionalmanner such as with igniter squibs. After the combustion of thisinnermost propellant layer is complete, a liquid oxidizer which ishypergolic with the liner material is injected or sprayed into thecombustion chamber. Contact of the liner with the oxidizer results inignition of the liner. Combustion of the liner is maintained bycontinuing the injection of oxidizer until the liner is consumed and thenext layer of propellant has been ignited. The ignition of the nextpropellant layer will result from contact of the layer with thehypergolic oxidizer and/or from the combustion of the liner. However,separate ignition means to ignite each layer of propellant can beprovided if desired. As many layers of propellant as desired can beignited and burned by this method of operating a rocket engine therebyaffording control over the total impulse produced by a given engine.

The possible schemes for contacting the fluid oxidizer and the linermaterial during operation of the engine are extensive. The simplestmethod is to continuously inject the oxidizer into the motor at alltimes during operation. Since the time to burn through a given number ofliners and propellant layers in a given motor is easily determinedaccording to known procedures, the total number of propellant layers tobe burned can be controlled by ceasing the fluid oxidizer flow at aprearranged time by means of a time-responsive cut-off valve. After theoxidizer flow is halted, no additional liners will be burned throughand, consequently, only the propellant layer burning at the time theoxidizer flow is stopped continues to burn.

Another means of controlling the omdizer flow is by means of a pressureresponsive valve. As is well known, the chamber pressure of a givensolid-fueled rocket engine produces a relatively uniform pressure curvefrom ignition to burn-out. Therefore,,the valve can be regulated so thatit will close when the chamber pressure falls to a certain predeterminedpressure and, thus, prevent the ignition and combustion of anyadditional liners. This in turn prevents combustion of additionalpropellant layers.

Still another means for regulating the liquid oxidizer flow is to use apressure-responsive valve which would open and admit the oxidizer whenthe chamber pressure began to suddenly drop as a result of the burn-outof one layer of propellant. The oxidizer would flow into the chamberigniting the liner and then the next propellant layer. This wouldincrease the chamber pressure and the valve would close, stopping theflow of oxidizer. The valve mechanism would be such that it would onlyopen and shut a predetermined number of times so that only the desirednumber of propellant layers would be burned.

The valve controlling oxidizer flow could also be actuated by means ofan accelerometer. Thus when a propellant layer burned out accelerationwould stop. At this point, the accelerometer would cause the valve toopen and thereby ignite the next propellant layer. A timing device wouldprevent further operation of the valve after a preset time interval hadlapsed.

The particular means for controlling the oxidizer flow is not criticalto the invention since many satisfactory art-recognized expedients areknown for regulating fluid flow. The selection of any given device foruse with a given motor is within the skill of the art. The onlyessential feature is that some means be provided to admit the oxidizerinto the chamber and to halt the oxidizer flow when the desired numberof propellant layers have been burned. The means of getting the oxidizerinto the combustion chamber can be any of the conventional methods nowused in the operation of liquid-fueled engines. Pumps and compressesgases are examples of suitable means for forcing the oxidizer into thecombustion chamber.

One embodiment of the present invention is illustrated in FIG. 1 andFIG. 2. In FIG. 2 a propellant grain l0 ofa type envisioned by thepresent invention is shown disposed within the wall 12 of the motor 30.The grain is made of three layers of propellant designated as l4, l6,and 18. The propellant layers are separated from each other by liners 20and 22. Four-point star perforation 24 is disposed along the center axisof the grain and runs from end to end therethrough. FIG. ll showspropellant grain l0 rigidly disposed in chamber 15 of motor 30. Thischamber is formed by the substantially cylindrical wall 12 of motor 30.The wall tapers at the rear end of the chamber to formconverging-diverging nozzle 70. The forward end 72 of the chamber isclosed except for opening 74. Opening 74 is in communication withperforation 24 which receives tube 48. A tank 40 is provided for storingthe liquid oxidizer. The oxidizer is transported from the tank to theforward portion 46 of perforation 2 3 through tube 48 which is providedwith nozzle 4 3. Oxidizer flow through tube 48 is regulated by valve 42.

In the operation of motor 30 according to one aspect of the presentinvention, propellant layer 18 is ignited and burned. Then a fluidoxidizer which is hypergolic with liner 20 is introduced into the motorchamber through nozzle 44. The oxidizer ignites liner 20 which burnsthrough and thereby ignites propellant layer 16. This process isrepeated until as many layers of propellant as required have been burnedto provide the necessary thrust for a specific task. The flow ofhypergolic oxidizer can be continuous and terminated only after thedesired number of propellant liners have been ignited.

FIG. 3 is a sectional view of another propellant grain configurationembodying the aspects of the present invention. The cylindrical grain 50is shown disposed in rocket motor chamber 52 formed by the cylindricalwall 82. Each of the propellant layers 54, 58 and 64 are separated fromeach other by liners 56 and 62. Cylindrical perforation 60 is disposedin the center of grain 50 and extends from end to end therethro'ugh.Operation of a rocket motor with a grain of this type would besubstantially identical to the operation of motor 30 which is discussedhereinabove.

As previously mentioned, the possible combinations of solid propellant,liner, and oxidizer are extensive. For example, the layers of propellantcan be of the polysulfide-perchlorate type shown in U.S. Pat. No.2,997,376. The liner material can be polybutadiene and the liquidoxidizer a perchloryl fluoridechlorine trifluoride solution containingabout 25 percent to about 50 percent by weight perchloryl fluoride.Polystyrene can also serve as the liner and fluorine as the oxidizer.These are but two specific examples of satisfactory combinations and areby no means limiting. Those skilled in the art are aware of the manyorganic polymers (practically all) which are hypergolic with theoxidizers mentioned hereinabove and will experience no difficulty infinding numerous satisfactory combinations of polymeric liners andhypergolic oxidizers within the framework and guidance of thisdisclosure.

No undue limitations should be attributed to the instant invention as aresult of the above detailed description thereof except as reflected inthe appended claims.

We claim:

1. A solid-fueled rocket propellant grain, said grain comprising incombination:

a. a plurality of concentric layers of propellant, the propellant layernearest the center of the grain having disposed therein a perforationpassing from end to end therethrough along the center axis of saidgrain, said propellant layers consisting essentially of a cured,intimate mixture of a polymeric binder, fuel particles, and an oxidizertherefor; and

b. a liner disposed between each of said layers in rigid relationshiptherewith to completely separate the adjacent surfaces of said layersfrom each other, said liner consisting essentially of a cured polymericmaterial incapable of supporting its own combustion.

2. The solid rocket propellant grain according to claim 1 wherein saidliner is up to about 0.75 inch in thickness.

3. A solid rocket propellant grain according to claim 2 wherein saidliner is a member selected from the group consisting of the polydienes,the polysulfides, the polyepoxides, polyurethanes, polylactams,polystyrenes, polyethylenes, polypropylenes, and polybutylenes.

4. A solid-fueled rocket motor, said motor comprising in combination:

a. a substantially cylindrical outside wall which forms a chamber, saidwall being tapered at the rearward end of said chamber to form aconverging-diverging nozzle, said wall also closing the forward end ofchamber except for an opening in communication with said chamber toreceive a tube whereby fluid can be introduced into said chamber; and

b. a solid rocket propellant grain rigidly disposed in said chamber,said grain comprising a plurality of concentric layers of propellant,the propellant layer nearest the center of the grain having disposedtherein a perforation passing from end to end therethrough along thecenter axis of said grain, said propellant layers consisting essentiallyof a cured intimate mixture of a polymeric binder,

fuel particles, and an oxidizer therefor, said layers being completelyseparated from each other by a liner, said liner consisting essentiallyof a cured polymeric material incapable of supporting its owncombustion.

5. The solid-fueled rocket motor according to claim 4 wherein said lineris up to about 0.75 inch in thickness.

6. The solid-fueled rocket motor according to claim 5 wherein said lineris a member selected from the group consisting of the polydienes, thepolysulfides, the polyepoxides, polyurethanes, polylactams,polystyrenes, polyethylenes, polypropylenes, and polybutylenes.

7. The method of operating a solid-fueled rocket motor to vary the totalimpulse achieved, said motor comprising, in combination, a substantiallycylindrical outside wall which forms a chamber, said wall being taperedat the rearward end of said chamber to form a converging-divergingnozzle, said wall also closing the forward end of said chamber exceptfor an opening to receive a tube whereby fluid can be introduced intosaid chamber; a solid rocket propellant grain rigidly disposed in saidchamber, said grain comprising a plurality of concentric layers ofpropellant, the propellant layer nearest the center of the grain havingdisposed therein a perforation passing from end to end therethroughalong the center axis of said grain, said layers consisting essentiallyof a cured intimate mixture of polymeric binder, fuel particles, and anoxidizer therefor, said layers being completely separated from eachother by a liner, said liner consisting essentially of a cured polymericmaterial incapable of supporting its own combustion, said methodcomprising:

a. igniting and burning the innermost layer of propellant,

thereby exposing the adjacent liner;

b. introducing into the forward end of said chamber through said openingan oxidizer fluid which is hypergolic with said liner whereby the linerignites, burns, and in turn ignites the adjacent layer of propellant;and

c. repeating steps (a) and (b) until as many propellant layers asdesired have been consumed.

8. The method according to claim 7 wherein said liner is up to about0.75 inch in thickness.

9. The method according to claim 8 wherein said liner is a memberselected from the group consisting of the polydienes, the polysulfides,the polyepoxides, polyurethanes, polylactams, polystyrenes,polyethylenes, polypropylenes, and polybutylenes.

10. The method according to claim 9 wherein said fluid oxidizer is amember selected from the group consisting of red fuming nitric acid,fluorine, chlorine trifluoride, perhalogenyl fluoride-halogen fluoridemixtures, perhalogenyl fluoridefluorine mixtures, and perhalogenylfluoride-chloryl fluoride mixtures.

1. A solid-fueled rocket propellant grain, said grain comprising incombination: a. a plurality of concentric layers of propellant, thepropellant layer nearest the center of the grain having disposed thereina perforation passing from end to end therethrough along the center axisof said grain, said propellant layers consisting essentially of a cured,intimate mixture of a polymeric binder, fuel particles, and an oxidizertherefor; and b. a liner disposed between each of said layers in rigidrelationship therewith to completely separate the adjacent surfaces ofsaid layers from each other, said liner consisting essentially of acured polymeric material incapable of supporting its own combustion. 2.The solid rocket propellant grain according to claim 1 wherein saidliner is up to about 0.75 inch in thickness.
 3. A solid rocketpropellant grain according to claim 2 wherein said liner is a memberselected from the group consisting of the polydienes, the polysulfides,the polyepoxides, polyurethanes, polylactams, polystyrenes,polyethylenes, polypropylenes, and polybutylenes.
 4. A solid-fueledrocket motor, said motor comprising in combination: a. a substantiallycylindrical outside wall which forms a chamber, said wall being taperedat the rearward end of said chamber to form a converging-divergingnozzle, said wall also closing the forward end of chamber except for anopening in communication with said chamber to receive a tube wherebyfluid can be introduced into said chamber; and b. a solid rocketpropellant grain rigidly disposed in said chamber, said grain comprisinga plurality of concentric layers of propellant, the propellant layernearest the center of the grain having disposed therein a perforationpassing from end to end therethrough along the center axis of saidgrain, said propellant layers consisting essentially of a cured intimatemixture of a polymeric binder, fuel particles, and an oxidizer therefor,said layers being completely separated from each other by a liner, saidliner consisting essentially of a cured polymeric material incapable ofsupporting its own combustion.
 5. The solid-fueled rocket motoraccording to claim 4 wherein said liner is up to about 0.75 inch inthickness.
 6. The solid-fueled rocket motor according to claim 5 whereinsaid liner is a member selected from the group consisting of thepolydienes, the polysulfides, the polyepoxides, polyurethanes,polylactams, polystyrenes, polyethylenes, polypropylenes, andpolybutylenes.
 7. THE METHOD OF OPERATING A SOLID-FUELED ROCKET MOTOR TOVARY THE TOTAL IMPULSE ACHIEVED, SAID MOTOR COMPRISING, IN COMBINATION,A SUBSTANTIALLY CYLINDRICAL OUTSIDE WALL WHICH FORMS A CHAMBER, SAIDWALL BEING TAPERED AT THE REARWARD END OF SAID CHAMBER TO FORM ACONVERGING-DIVERGING NOZZLE, SAID WALL ALSO CLOSING THE FORWARD END OFSAID CHAMBER EXCEPT FOR AN OPENING TO RECEIVE A TUBE WHEREBY FLUID CANBE INTRODUCED INTO SAID CHAMBER; A SOLID ROCKET PROPELLANT GRAIN RIGIDLYDISPOSED IN SAID CHAMBER, SAID GRAIN COMPRISING A PLURALITY OFCONCENTRIC LAYERS OF PROPELLANT LAYER NEAREST THE CENTER OF THE GRAINHAVING DISPOSED THEREIN A PERFOROATION PASSING FROM END TO ENDTHERETHROUGH ALONG THE CENTER AXIS OF SAID GRAIN, SAID LAYERS CONSISTINGESSENTIALLY OF A CURED INTIMATE MIXTURE OF POLYMERIC BINDER, FUELPARTICLES, AND AN OXIDIZER THEREOF, SAID LAYERS BEING COMPLETELYSEPARATED FROM EACH OTHER BY A LINER, SAID LINER CONSISTING ESSENTIALLYOF A CURED POLYMERIC MATERIAL INCAPABLE OF SUPPORTING ITS OWNCOMBUSTION, SAID METHOD COMPRISING: A. IGNITING AND BURING THE INNERMOSTLAYER OF PROPELLANT, THEEBY EXPOSING THE ADJACENT LINER; B. INTRODUCINGINTO THE FORWARD END OF SAID CHAMBER THROUGH SAID OPENING AN OXIDIZERFLUID WHICH IS HYPERGOLIC WITH SAID LINER WHEREBY THE LINER IGNITES,BURNS, AND IN TURN IGNITES THE ADJACENT LAYER OF PROPELLANT; AND C.REPEATING STEPS (A) AND (B) UNTIL AS MANY PROPELLANT LAYERS AS DESIREDHAVE BEEN CONSUMED.
 8. The method according to claim 7 wherein saidliner is up to about 0.75 inch in thickness.
 9. The method according toclaim 8 wherein said liner is a member selected from the groupconsisting of the polydienes, the polysulfides, the polyepoxides,polyurethanes, polylactams, polystyrenes, polyethylenes, polypropylenes,and polybutylenes.
 10. The method according to claim 9 wherein saidfluid oxidizer is a member selected from the group consisting of redfuming nitric acid, fluorine, chlorine trifluoride, perhalogenylfluoride-halogen fluoride mixtures, perhalogenyl fluoride-fluorinemixtures, and perhalogenyl fluoride-chloryl fluoride mixtures.